1. Field of the Invention
The present invention relates to a propulsion thrust control system configured for control of a rocket-propelled vehicle.
2. State of the Art
Solid propellant rocket motors employ a propellant comprising a solid fuel charge or “grain” which burns to generate exhaust gases and other combustion products, which are expelled through one or more nozzles of the rocket motor to provide thrust. Once a grain of solid propellant is ignited, it is difficult to extinguish and the entire grain is ordinarily consumed after ignition. Additionally, effecting variation of thrust is more difficult in solid propellant than in liquid propellant rocket engines. However, simple structural design of solid propellant rocket motors and ease of storage of the solid propellant are advantages of the solid propellant motor.
Attitude control, in the form of influencing the pitch, yaw, and/or roll of the rocket assembly in flight, may be accomplished with a thrust vector control (TVC) system or a separate attitude control system (ACS). A TVC system may comprise an axial thrust nozzle rotationally positionable at a desired angle within a range offset from the longitudinal axis of the rocket motor to alter the vector at which the combustion products exit the rocket motor. Repositioning of the nozzle alters the direction of the forces acting on the vehicle in which the rocket motor is installed to alter the vehicle's direction of flight. Single, movable TVC nozzles provide adequate control over the rocket assembly's yaw and pitch, but do not provide any significant degree of roll control.
Multiple rocket engines or gas generators and associated thrusters are often employed to control attitude. The rocket engines or thrusters are offset from the longitudinal axis of the rocket motor assembly so that firing of selected ones or groups of the engines or powering of selected ones or groups of thrusters enables attitude control over the rocket motor assembly. Use of a separate ACS in combination with one or more axial thrust engines or thrusters increases the weight of the rocket motor assembly due to the additional hardware. A separate ACS may use a solid-propellant gas generator directly connected to a manifold providing a selective hot gas flow to nozzle valve clusters. Roll control may be provided by the ACS or through the inclusion of a separate roll control system (RCS). Separate gas generators and thrusters may be provided for the RCS.
As stated, each of the control systems such as the ACS and the TVC direct gases through valves that results in the generation of thrust for altering the vehicle's direction of flight. Valves are known that may be configured in either “open” or “closed” states. Additionally, valves that include at least a third or “partially open” state are known as “proportional valves.” The state of each valve is determined by the cross-sectional area of the orifice or “throat” for each valve. The state of each of the valves is controlled by a motor controller or control system, which adjusts the state of the various valves based upon one or more control inputs.
One method for controlling the state of each of the valves relies upon pressure measurements which become the control inputs or process variables to the control system. The control inputs in this mode of control are net thrust per valve set (Fnet,i) and a pressure set point or limit. As used herein, net thrust is defined as the differential thrust between opposing valve pairs.
According to a thrust/pressure-only control methodology, the total throat area is regulated to obtain a pressure. An approximate throat area (control variable) for a pressure command is given in the following equation:At=(C*ρAsr(Pcmd/Pref)n)/(gcPcmd)  [1]                where: ρ=density of propellant [lbm/in3]                    Pref=reference pressure used to determine r [lbf/in2]            As=surface area of propellant [in2]            n=exponent            r=burn rate @ Pref [in/sec]            At,i=throat flow area of valve i [in2]            Pcmd=gas generator pressure command [lbf/in2]            gc=gravitational constant [(lbm/lbf)(ft/s2)]            C*=characteristic exhaust velocity [ft/sec]where the pressure command could be generated by the following equation:Pcmd=max {C1max(Fnet,i), C2 max(Fnet,i)1/n}  [2]                        where: C1=1/(At(FULL OPEN) CF)                    C2=Pref {1/(ISP ρ AS r)}1/n                         where: ISP=specific impulse [s]                    CF=discharge coefficient of valve                        
The pressure command is determined by taking the maximum of two calculations. The first involves the pressure required to meet the worst case (highest) net thrust command relative to the available throat area for a given valve set. This pressure is determined by the specific valve characteristics. The second involves the pressure required to meet all net thrust requirements. This pressure is determined by propellant characteristics and geometry,
The net thrust from a given valve set is proportional to the gas generator (GG) pressure and difference in regulated throat areas of opposing valves:Fnet,i=(At, i−At,j)P CF, where valves i and j comprise a valve set and provide thrust in opposite directions.  [3]
The distribution of the regulated throat area meets thrust commands in accordance with the system requirements. By way of simplified example and to avoid infinite solutions of the pressure-only control methodology, the throat areas are evenly distributed (in proportion to flow capacities) among all valves and the same offset, in opposite directions, is added to each valve (in a valve set) to achieve the net thrust command. It is known that the faster response times of the valves compared to slower response of the gas generator, allow controlling net thrust during transient events in the gas generator. This is especially true, for example, when controlling ACS valves that are much smaller than divert valves since ACS valves have a much smaller effect on gas generator (GG) pressure.
While the thrust control approach using a pressure-only methodology may be effective, inefficiencies remain, specifically, the sensitivity to the GG variation. For example, if the propellant is not burning at the specifically designed burn rate or propellant surface area deviates from that designed, a differential (i.e., more or less) mass flow results. For example, assume that more mass is generated than desired because the actual burn rate is higher than designed. If pressure is regulated, the valves will have to open more than predicted to accommodate the larger actual mass flow generated per unit time (“m-dotgen”), resulting in an excess (i.e., wasted) thrust (i.e., more mass flow discharged per unit time (“m-dotdisch”) than required to meet all the net thrust commands) as shown by:Σm-dotdisch=P ΣAt,igc/C*>(ΣFnet,i)/Isp=m-dotcmd.  [4]
As an example, demand for thrust in a specific direction occurs by commanding or controlling a valve pair to achieve the commanded thrust. However, the aggregate throat area of all the valves must compensate for pressure regulation, while each opposing pair has the correct differential throat area. Inherently, the pressure compensation requires opening more than predicted resulting in wasted propellant by dumping some of the generated gas. Propulsion systems that incorporate solid propellant, while desirable due to performance and weight reduction, become less desirable as inefficiencies are introduced due to dumping of gases that could otherwise be used for motion control.
For thrust control using a pressure-only methodology, an updated burn rate coefficient (rupdate) can be calculated whenever the system is in a quasi-steady-state condition, as defined when the pressure is within a defined percentage of the commanded value and oscillates below a threshold value. In this condition, an updated burn rate coefficient may be simply calculated by knowing measured gas generator (GG) pressure (Pmeasured) and measured total throat areas (At, measured, i) achieving that pressure with all system parameters assumed constant as shown by:rupdate=ΣAt,measured, i gc Pmeasured/(C*ρAs(Pmeasured/Pref)n).  [5]
Knowledge of the actual burn rate coefficient helps to predict propellant burn-out, but does not directly aid in thrust control when implementing a thrust control methodology based on pressure-only control inputs. Additionally, while a pressure command change (ΔP (=Pnom−Pupdate) can be subtracted from the pressure command:Pnom=(As ρrnomC*/(ΣAt,nomPnrefgc))(1/(1−n))  [6]Pupdate=(AsρrupdateC*(ΣAt,nomPnrefgc))(1/(1−n))  [7]Pcmd=Pcmd(Pnom−Pupdate)  [8]to compensate for the burn rate variability with Pnom representing the pressure at a nominal burn rate coefficient and nominal total throat area and Pupdate representing the pressure at the updated burn rate and nominal total throat area, an updated ΔP must be calculated any time a new estimate of the burn rate coefficient occurs or any time the pressure command changes.
In view of the above-enumerated deficiencies in the state of the art with respect to pressure-only thrust control of a rocket-propelled vehicle, it would be desirable to develop a methodology for controlling thrust in a vehicle for improving the inefficiencies and for calculating thrust control commands.